A solar array, as defined herein, pertains to a structure which is stowable in a small volume for shipment and launch, and that is deployable when in space to expose a large surface area of photovoltaic collectors (solar cells) to the sun, and that is attached to certain spacecraft vehicles, to provide power for spacecraft operations. FIG. 1 shows a typical spacecraft (101) that uses a solar array (102) for power production, with the solar array (102) in the deployed configuration.
Solar arrays typically consist of an underlying structure for deployment of a substantial number of individual photovoltaic solar cells from the body of a spacecraft. Once fully deployed, it is desirable for the solar array structure to provide a lightweight, stiff, strong, stable and flat platform for the solar cells to allow uniform exposure to the sun and minimize on-orbit spacecraft attitude control disturbance loads. Solar arrays are typically stowed such that they are constrained in a small envelope prior to and during launch of the spacecraft and then are deployed to their fully extended configuration, exposing the maximum area of solar cells once the spacecraft has reached its position in outer space. It is desirable to minimize the volume of the stowed package while at the same time maximizing the available solar cell area that can be packaged when stowed, and subsequently deployed to allow for maximum power production for the spacecraft.
In certain prior art applications of solar arrays, the structure consists of flat rigid panel substrates that are configured for stowage by means such as hinges which will permit the panels to be folded against each other to minimize the dimensions of the array in the stowed configuration. Folding of rigid panels involves mechanical items such as hinges and latches; and actuating mechanisms such as springs, cables and pulleys which must be highly reliable to prevent complete loss of a spacecraft and its payload due to inability to deploy the power-producing array. These mechanical components are costly, and involve added weight which is desirable to minimize. An example of such an array is shown in: Everman et al U.S. Pat. No. 5,487,791.
In order to allow for further reduction in the deployable solar arrays weight and stowed volume, the solar cell mounting can be configured using a flexible substrate, or blanket. Various flexible solar cell blanket substrates have been used, such as those fabricated from a fiberglass mesh or thin polymeric sheet upon which are bonded the numerous crystalline solar cells. Flexible-blanket solar arrays have been typically been limited to crystalline solar cell arrays packaged in a long roll or pleated stack that is deployed using a separate deployment boom or hub structure requiring external motor power for deployment motive force. These flexible array deployment structures have typically consisted of very complex mechanical systems such as coilable or articulated truss booms, or radially oriented spars that rotate about a central hub, which can add undesired parts, complexity, weight and cost to implement. Examples of prior art flexible blanket arrays are shown in the following United States patents: Harvey et al U.S. Pat. No. 5,296,044; Stribling et al U.S. Pat. No. 6,983,914; Hanak et al U.S. Pat. No. 4,636,579 and Beidleman et al U.S. Pat. No. 7,806,370.
Critical to ensuring deployment reliability is to allow for maximum deployment motive force (or torque) in the design of the deployment actuators. Reliability is enhanced when the deployment actuation has a large force margin (typically required to be at least 3:1) over any and all predicted (and unforeseen) sources of resistance to deployment, such as harness bending, friction in joints, snagging or adhesive sticking between blanket layers. Historical solutions used to increase deployment force margin in a linearly deploying boom have been to increase the size and capability of structural components and use deployment actuators (such as springs or motors) to “force” the boom out, further increasing weight and complexity. An example of this sort of prior art is the bi-stem booms used on the Hubble telescope solar array to unfurl the solar blankets. In this application, the booms are comprised of paired curled sheets of metal that are rolled and nest within one other to form a cylindrical boom upon deployment, and a complex and heavy motorized mechanism is used to externally push the boom material out in a known deployment direction and with sufficient force. Utilizing the elastic strain energy inherent in the deployment boom material alone to achieve high deployment force has not been successfully used in the prior art for deployable boom-type solar array structures. This is because in order to raise the available actuation energy to levels sufficient to achieve acceptable deployment force margin, typical metallic or fiber-composite materials are too highly stressed (they are unacceptably close to failure), and the kinematics of deployment are difficult to control and predict due to the high internal energy and ungoverned nature of the stowed assembly upon release.
It is also desirable to maximize the deployed natural frequency (stiffness) and strength (against deployed accelerations) of a solar array. As the size of the solar cell deployed area and the solar array supporting structure increase, the stiffness of the solar cell array decreases and, as a result, the vibration frequency decreases and disturbance deflections increase. The ability of the spacecraft attitude control system to orient the spacecraft may be impaired if the deflections due to low-frequency solar array movement are excessive.
A review of the prior art shows that significant efforts have been made to reduce the weight and increase the deployment reliability and force margins of rigid and flexible blanket solar arrays for a given set of deployed stiffness and strength requirements. These prior efforts have resulted in solar array designs tending to involve difficult and time consuming manufacturing, higher complexity and higher cost.
The current larger market for spacecraft is demanding significant decreases in the cost of all spacecraft systems and payloads, including solar arrays. As the demand for spacecraft power grows, it is desirable to provide a deployable solar array system that permits straightforward scaling up in size to allow use of larger deployed solar cell areas. It is also desirable to enhance reliability, while at the same time reducing weight and cost, by reducing the number of component parts and mechanisms required to achieve deployment and adequate deployed performance. Because mechanical components are subject to failure, and must be rigorously tested as an assembled system to validate their reliability; solar array reliability can be increased significantly, while simultaneously reducing cost and mass, by reducing the amount of mechanical components and mechanisms required to deploy and form the array into a deployed structure.